1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a blade outer air seal and its cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In an axial flow gas turbine engine, a compressor provides compressed air to a combustor where a fuel is mixed with the compressed air to produce an extremely high temperature gas flow, the resulting hot gas flow being passed through a multiple stage turbine to produce power that drives the turbine rotor shaft. In an aero engine used to power an aircraft, the turbine is used mainly to drive the compressor and an optional fan blade to propel the aircraft. In an industrial gas turbine engine, as much of the hot gas flow is used to drive the turbine in order to convert as much of the chemical energy from the combustion into mechanical work that is used to drive an electric generator.
The turbine section of the engine includes a plurality of stages of rotor blades that convert the hot gas flow into mechanical energy that drives the turbine shaft. The rotor blades rotate within the engine and form a gap between the blade tip and an outer shroud on the engine casing. The blade tip gap will allow for hot gas flow to leak through the turbine, and therefore a loss of energy is produced. The blade tip gap will change depending upon the outer shroud and blade temperatures. Limiting the gap space such that a hot gas flow leakage is minimized will improve the efficiency of the engine.
In the prior art, blade outer air seals (or, BOAS) have been proposed to limit the blade tip gap formed between the blade tip and the outer shroud. The outer shroud seal is formed from a plurality of arcuate shroud segments forming an annular arranged on the stator assembly that encircles the rotor blades. Thermal barrier coatings (TBC) have also been added to the shroud segment surfaces to limit thermal damage to the BOAS since the hot gas flow leakage will affect the blade tip and the shroud segment material.
One problem found with arcuate shroud segments that have a TBC applied thereto is the stresses developed between the TBC and the substrate on which the TBC is applied. When the engine is cold, no stress is developed between the TBC and the substrate of the arcuate shroud segment because the TBC was applied in the cold condition. When the engine is operating and the shroud segments reach normal operating temperatures, the shroud segments tend to bend due to thermal growth. Mismatch between the coefficients of thermal expansion between the TBC and the shroud segment will induce high stresses between the materials and cause spalling of the TBC.
One prior art reference that attempts to address this problem is U.S. Pat. No. 5,375,973 issued to Sloop et al on Dec. 27, 1994 entitled TURBINE BLADE OUTER AIR SEAL WITH OPTIMIZED COOLING in which the BOAS shroud segment includes first and second groups of cooling passages each with a cooling air supply orifice to supply cooling air from the casing cavity, one or more re-supply holes connecting the cooling passages to the cavity to re-supply cooling air, and cross supply orifices connecting adjacent cooling passages to provide for cross flow of cooling air between the cooling passages in the event that the flow within a particular passage shroud decrease, as in the case where the metering orifice of a cooling passage is partially obstructed by a foreign object (see column 6, lines 25-34 in the Sloop et al patent). The Sloop et al patent provides improved BOAS cooling over the cited prior art. However, the present invention improves over the Sloop et al cooling design by providing for multi-metering diffusion compartment cooling to provide improved cooling using less cooling air.